1. Field of the Invention
The invention relates generally to gas turbine engines, and more particularly to controlling the radial clearance between a turbine rotor blade tip and a stator shroud assembly.
2. Description of Related Prior Art
In a turbine engine, combustion gases pass across rotatable turbine blades to convert the energy associated with combustion gases into mechanical motion. A shroud assembly tightly encircles the turbine blades to ensure that combustion gases are forced over the turbine blades and do not pass radially around the turbine blades. It is desirable to maintain the smallest possible gap between the tips of the turbine blades and the shroud assembly to maximize the efficiency of the turbine engine. However, a challenge in maintaining the smallest possible gap arises because the turbine blades can expand radially during various phases of engine operation at a rate that is much greater than a rate at which the shroud assembly can radially expand. For example, when the power output of the turbine engine rapidly increases, such as during take-off in a turbine used for aircraft propulsion, the turbine blades will increase in radial length rapidly and the tips of the turbine blades may penetrate the inner linings of the shroud assembly. This could damage both the turbine blades and the shroud assembly. Also, this event can compromise the capacity of the shroud assembly to maintain the smallest possible gap during periods of relatively low power production.